1. Field of the Invention
The present invention relates to high speed gun or artillery launched projectiles, and more particularly to airbreathing propulsion-assisted projectiles which accelerate after launch using a combination of ramjet and scramjet propulsion to enable flight at hypersonic velocities.
2. Description of the Related Art
The ramjet and supersonic ramjet (scramjet) propulsion cycles for supersonic and hypersonic (Mach &gt;5) engines are well-known within the art of aerospace propulsion. For the sake of discussion here, consider an engine defined by an external compression device or forebody, an internal compression device such as an inlet including a diffuser and an isolator, a combustion device or combustor, and an expansion device or nozzle. All surfaces wetted by flow streamlines ultimately passing through the engine are considered to be a part of the engine flowpath since they contribute to the engine cycle performance. Consequently, the integration of the airframe and the propulsion systems for vehicles or projectiles employing these propulsion cycles is critical for high performance.
In the ramjet propulsion cycle, high velocity air is compressed through a series of forebody and inlet shocks and through a subsonic diffuser, all of which decelerate the air to a subsonic velocity near the fuel flame speed. Fuel is injected into a combustor and conventional subsonic combustion occurs, thereby increasing the temperature and pressure of the flow. The high pressure gas is then expanded through a nozzle, increasing the velocity and momentum of the flow to produce thrust. This cycle is efficient for freestream Mach numbers ranging between approximately 2 and 5. However, for freestream velocities above about Mach 5, the temperatures and pressures associated with decelerating the flow to subsonic speeds for combustion are severe and begin eroding engine cycle performance.
The static temperature at the combustor entrance approaches the stagnation temperature and dramatically impacts fuel combustion. At such extreme temperatures, an appreciable amount of the energy which would be released due to combustion is bound in dissociated air and combustion product molecules such that the temperature rise due to combustion is reduced. The energy contained in dissociated gases is largely unavailable for the expansion and acceleration of the exhaust mixture and thrust is lost as a result.
For Mach numbers above 5, the main advantage of scramjet propulsion is that supersonic velocities within the combustion chamber are accompanied by lower static temperatures, pressures, and reduced total pressure losses. By reducing combustion product dissociation reduced temperatures increase combustion efficiency, reduced pressures decrease loads on engine structure, and reduced total pressure losses (entropy gains) increase the flow energy available for thrust production.
A large number of parameters impact the specific impulse (I.sub.sp, or thrust per pound of propellant) performance of ramjet and scramjet systems. They include, but are not limited to, the forebody and inlet contraction ratios, the inlet efficiency, the fuel mixing efficiency, the combustor efficiency, and the nozzle efficiency.
The purpose of the inlet is to capture a desired quantity of air flow and deliver it to the combustor at a desired pressure and Mach number with a minimum of entropy producing losses. The technology and parameters necessary to successfully design and operate an efficient supersonic inlet are well-known but difficult to capture in a single design. The mass flow captured by the inlet compared to the drag of the vehicle must be sufficiently large that a net thrust can be expected across the entire Mach number range of operation for achievable values of ramjet or scramjet I.sub.sp performance.
The isolator (also known as a constant area diffuser) is located between the inlet and the combustor entrance, and is necessary to adjust flow static pressure from that of the inlet exit to the higher combustor pressure downstream during ramjet and early scramjet ("dual-mode") operation. When combustor pressure rise is large and inlet Mach numbers low, as in ramjet operation, boundary layer separation in the combustor can lead to inlet interaction and engine unstart. An isolator permits a shock train to develop between the inlet and combustor with a near normal shock static pressure rise without any upstream inlet interaction. The length of the isolator is critical to carrying out this function.
The combustor provides the physical domain for injecting a liquid or gaseous fuel into high velocity air and mixing the fuel and air for combustion. The fluid and chemical phenomena present in the combustor are extremely complex and include the effects of laminar and turbulent mixing of fuel injection jets with boundary layers and core flows, and the finite rate chemical kinetics of the exothermic combustion reactions. Fuel ignition and flameholding are also important issues. Some of the typical design parameters are the fuel injection geometry, mixing enhancement devices, and the length of the combustor required to achieve the high mixing and combustion efficiencies necessary for high Isp performance across the Mach number range of interest. Fuel injection location and mixing rate (i.e., distribution of heat release) is also important for controlling if and where flow choking (Mach 1) occurs in the combustor. Fuel is generally injected aft in ramjet mode, both fore and aft in dual-mode (combined supersonic and subsonic combustion), and forward in scramjet mode.
The nozzle or expansion system is important to the specific impulse Isp of the projectile because it produces thrust by accelerating the high static pressure flow exiting the combustor to lower pressure and high velocity (i.e., momentum). Typically composed of internal and external nozzles, the objective is to expand the high pressure flow to the lowest pressure possible using a shape that minimizes the combination of friction losses, chemical recombination losses, and flow divergence (angularity) losses.
The ratio of the nozzle expansion area to the inlet capture area and the ratio of inlet mass flow to nonflowpath drag are critical figures of merit in designing a system which produces a flowpath thrust that exceeds the nonflowpath drag and therefore produces acceleration. Balancing the geometric details for high efficiencies and high Isp performance with the vehicle drag is the traditional challenge inherent in ramjet and scramjet vehicle design.
Methods to improve the range and velocity performance of gun or artillery launched projectiles have been investigated since the earliest development of these devices. Gun projectile velocities are ultimately limited by the speed of sound in burned propellant gases. Therefore, methods for accelerating the projectile after it leaves the barrel have also been investigated.
One approach successfully developed and commonly used today is that of a rocket-assisted projectile. A rocket motor containing a solid or liquid propellant is incorporated into the base of the projectile and ignited after leaving the barrel. The range and acceleration potential of these rocket assisted projectiles is rather limited due to the small amount of fuel which can be carried and the relatively low I.sub.sp which can be produced by a rocket motor in comparison to an air breathing ramjet or scramjet propulsion cycle.
Research in supersonic air-breathing propulsion systems for aircraft and missiles has been in progress since the 1940's. As empirical knowledge grew in the late 1950's, researchers investigated propulsion for hypersonic aircraft and missiles using scramjet engines. Research into scramjet propulsion continued during the 1970's at the NASA Langley Research center, and in the 1980's and 1990's grew considerably under the auspices of the National Aerospace Plane program. Unfortunately, no scramjet engines have been demonstrated outside of a wind tunnel. Again, these research activities focused exclusively on aircraft and missile applications.
U.S. Pat. No. 4,291,533 to Dugger et al. describes a rocket-launched scramjet powered missile. Typical gun-launched projectile accelerations of tens of thousands times greater than the acceleration of gravity (g) far exceed the approximately hundred g acceleration of a typical high performance rocket booster, and consequently would prevent directly adapting the missile design described or any other missile to gun launching to the velocity required to initiate a ramjet or scramjet engine.
Application of ramjet propulsion to gun-launched projectiles have been described by Olson et al. in U.S. Pat. No. 5,067,406 and by Flatau in U.S. Pat. No. 4,539,911. These patents considered tubular projectiles utilizing a solid propellant. These concepts focused on producing thrust which essentially balanced the aerodynamic drag in order to reduce the deceleration of the projectile as opposed to accelerating beyond the muzzle velocity. The lightweight construction of these projectiles also severely constrains the gun launch acceleration loads which the projectiles can survive and restricts the muzzle launch velocity. The flowpath through the center of the projectile also limits the fuel which can be carried and hence the projectile range which can be achieved or increase in velocity if thrust exceeds aerodynamic drag. These concepts cannot carry a payload of significant volume or size due to interference with the propulsive flowpath.
Botwin et al. in US Pat. No. 4,428,293 addresses the payload and fuel volume issues of the previously referenced patents to Olson et al. and Flatau. However, Botwin specifically discloses that the ramjet powered projectile is designed to maintain a thrust-to-drag balance such that it follows a predetermined vacuum ballistic trajectory.
A scramjet system launched from a light gas gun for scramjet propulsion testing and experiments in a closed test chamber was documented in 1968 by H. H. King and O. P. Prachar in the Air Force Aero Propulsion Laboratory Technical Report AFAPL-TR-68-9. This study represents the only known attempt to launch a scramjet-shaped projectile from a gun barrel, and was conducted to investigate issues pertaining to launch and acceptable free flight of an annular combustor scramjet model. The scramjet model was too small to include a fuel system, and was therefore limited to unfueled launches to verify structural integrity and aerodynamic stability. Fuel systems were tested separately in simple cones only, not in scramjets.
The design of the model included an internal contraction ratio (i.e., the ratio of the inlet area at the cowl leading edge to the minimum flow area downstream of the cowl leading edge) of unity for positive inlet starting characteristics, but with a very low airflow capture area to drag ratio with the result that a net thrust or positive acceleration could not be produced even if it was fueled.